There is a rapidly growing commercial demand for small satellites, and, more particularly, for satellite instruments that demand precision control of position and orientation on small and large spacecraft. However, there is concomitant need for a significant reduction in the cost and complexity of these types of spacecraft and all of their subsystems for there to be broad application of such satellites in the commercial space industry.
In conventional satellites, thrust for station control and rapid maneuvers is provided to a spacecraft by chemical propulsion, such as via hydrazine or other rocket motors. However, the exhaust velocity of such chemical rockets is limited by the inherent specific energy released by combustion, accordingly, chemical rockets burn up more propellant to effect an orbital maneuver than would other forms of propulsion. Furthermore, the propellant reservoirs and feed systems for these types of chemical propulsion add prohibitively to the size, weight and complexity of the spacecraft, making them unusable for very small spacecraft or for highly distributed propulsion on very large space structures.
One form of propulsion that has gained a lot of recent interest is the electrospray thruster. An electrospray thruster is a form of electric propulsion for spacecraft that creates thrust from liquid propellants by ejecting and accelerating charged particles. Electric thrusters are categorized by how they accelerate the ions, using either electrostatic or electromagnetic forces. Electrostatic electrospray thrusters use the Coulomb force and accelerate charged particles in the direction of the electric field. Electromagnetic ion thrusters use the Lorentz force to accelerate the charged particles. Electrospray thrusters are more efficient than ion and Hall electrostatic thrusters. They also have the potential to be much more scalable in size, mass and thrust range to be applicable to very large and very small spacecraft. They currently operate at very low thrust levels on the order of microNewtons. The drawback of the low thrust is low spacecraft acceleration because the mass of current electric power units is directly correlated with the amount of power given. This low thrust makes electric thrusters unsuited for launching spacecraft into orbit, but they are ideal for in-space propulsion applications, such as station keeping, maneuvering and precision pointing.
Regardless of the method of accelerating the ions, all electrostatic thruster designs take advantage of the charge/mass ratio of the charged particles. Relatively small potential differences can create very high exhaust velocities that are on the order of thousands of seconds. As shown in FIG. 1A, a FEEP, and, for that matter any electrospray thruster, device consists essentially of an emitter, extractor electrode and possibly an accelerator electrode (downstream of the extractor electrode). (See, A. Genovese, et al., AIAA 2004-3620, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, July 2004, the disclosure of which is incorporated herein by reference.) The emitter is an externally wetted needle or an internally wetted capillary tube. A potential difference of the order of 5-10 kV is applied between the two electrodes, which generates a strong electric field at the tip of the emitter. The interplay of the electrostatic and surface tension forces generates surface instabilities, which give rise to Taylor cones (inset) on the liquid surface. At sufficiently high values of the applied field, charged particles are extracted from the cone tip by field evaporation or field ionization, which then are accelerated to high velocities on the order of tens of kilometers per second. A separate negative charged particle source is required to maintain spacecraft charge neutrality. This process of creating and accelerating charged particles is very efficient, with beam efficiencies greater than 90%.
Electrospray thrusters have been demonstrated and flight qualified to operate in either charged droplet at hundreds of seconds of specific impulse or in an ion emission mode at thousands of seconds at low thrust levels. (See, J. Ziemer, et al., AIAA 2008-4826, 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, July 2008, the disclosure of which is incorporated herein by reference.) The specific charge of the charged particles with some propellants in these thrusters could be controlled for extraordinary specific impulse and thrust range capability. This propulsion approach reduces the amount of reaction mass or fuel required, but increases the amount of specific power required compared to chemical propulsion. Due to their very low thrust level capability (in the microNewton range), electrospray thrusters are primarily used only for microradian, microNewton attitude control on spacecraft, such as in the ESA/NASA LISA Pathfinder scientific spacecraft.
A number of different “fuels” can be used in these electrospray thrusters. Ionic liquids are typically used in ‘colloid’ electrospray thrusters to generate charged droplet beams, and have been demonstrated in ion and droplet emission modes and in positive and negative charged particle emission modes. The ionic liquid propellant is typically pushed to the emitter tip through a capillary tube from a pressurized reservoir and controlled by a piezo valve. Field Emission Electric Propulsion (FEEP) electrospray thrusters typically use liquid metal (usually either cesium or indium) as a propellant and generate ion beams. The propellant is stored as a solid, melted to flow, and pulled to the emitter tip along external grooves by capillary forces. Both types of propellants have been demonstrated in each type of emitter and have been demonstrated in both ion and droplet emission modes for a unique specific impulse range capability among electric thrusters of several hundred to several thousand seconds.
Examples of conventional FEEP devices can be found in U.S. Pat. No. 4,328,667 to Valentian et al., which describes a liquid metal ion thruster assembly having a plurality of hollow-cone tips coupled to a reservoir of liquid metal, where the metal ions are drawn from the tip by the electrostatic force generated by an adjacent electrode; U.S. Pat. Nos. 6,097,139 and 6,741,025 to Tuck et al., which describe the use of impurities on a surface for the formation of enhanced electric fields for use as composite field emitters; U.S. Pat. No. 6,516,024 to Mojarradi et al and U.S. Pat. No. 6,996,972 to Song, which describe a hollow tip liquid ion extractor assembly for generation of thrust; U.S. Pat. No. 7,059,111, which describes a thruster system whereby liquid metal ions are boiled from a reservoir and electro-statically attracted through a cylindrical ring, thereby generating thrust; and U.S. Pat. Nos. 6,531,811 and 7,238,952, which both describe an ion extractor having a reservoir opposite a needle tip and an extractor electrode, the disclosures of each of which are incorporated herein by reference.
Although there are a large number of prior art on liquid metal ion propulsion devices, none of these devices address the fundamental issues limiting the scalablilty and applicability of this thruster technology, which is low thrust density. Addressing this issue with microfabricated thruster components, a common propellant reservoir and a capillary force driven feed system will simultaneously improve the thrust range, thrust density, system performance, mass, volume and cost by more than 10× to enable this technology for an extremely broad range of mission applications and revolutionary propulsion capabilities.